This disclosure relates to a gas turbine engine, and more particularly to a turbine vane that may be incorporated into a gas turbine engine. The vane's airfoils and platforms include a plurality of film cooling holes as part of a cooling arrangement.
Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.
Both the compressor and turbine sections may include alternating series of rotating blades and stationary vanes that extend into the core flow path of the gas turbine engine. For example, in the turbine section, turbine blades rotate and extract energy from the hot combustion gases that are communicated along the core flow path of the gas turbine engine. The turbine vanes, which generally do not rotate, guide the airflow and prepare it for the next set of blades.
Each of the blades and vanes include airfoils that extend into the core flow path of the gas turbine engine between inner and outer platforms. Cooling airflow is communicated into an internal core of the airfoil and can be discharged through the plurality of film cooling holes to provide a boundary layer of film cooling air along the external surface of the airfoil and platforms. The film cooling air provides a barrier that protects the underlying substrate of the vane from the hot combustion gases that are communicated within the core flow path. By way of example U.S. Published Patent Application No. 2015/0211376 describes a hole configuration for a turbine vane.